Gas turbine engine mid turbine frame with flow turning features

ABSTRACT

A gas turbine engine includes first and second stages having a rotational axis. A mid turbine frame is arranged axially between the first and second stages. The mid turbine frame includes a circumferential array of airfoils, and the airfoils each have a curvature provided equidistantly between pressure and suction sides. The airfoils extend from a leading edge to a trailing edge at a midspan plane along the airfoil. An angle is defined between first and second lines respectively tangent to the intersection of the plane and the curvature at airfoil leading and trailing edges. The angle is equal to or greater than about 10°, for example. In one example, an airfoil aspect ratio is less than 1.5.

This application claims priority to U.S. Provisional Application No.61/593,162, which was filed on Jan. 31, 2012.

BACKGROUND

This disclosure relates to a gas turbine engine mid turbine frame withflow turning features.

One typical gas turbine engine includes multiple, nested coaxial spools.A low pressure turbine is mounted to a first spool, and a high pressureturbine is mounted to a second spool. A mid turbine frame is arrangedaxially between the low pressure turbine and the high pressure turbine.One example mid turbine frame includes first and second circumferentialarrays of turbine vanes adjoining radially spaced outer and inner cases.The first and second array of vanes are axially spaced from one another.Oil and air may be passed through the airfoils.

SUMMARY

A gas turbine engine includes first and second stages having arotational axis. A mid turbine frame is arranged axially between thefirst and second stages. The mid turbine frame includes acircumferential array of airfoils, and the airfoils each have acurvature provided equidistantly between pressure and suction sides. Theairfoils extend from a leading edge to a trailing edge at a midspanplane along the airfoil. An angle is defined between first and secondlines respectively tangent to the intersection of the plane and thecurvature at airfoil leading and trailing edges. The angle is equal toor greater than about 10°.

In a further embodiment of any of the above, the midspan plane isoriented at a flow path angle relative to the rotational axis in therange of 20°-60°.

In a further embodiment of any of the above, the mid turbine frameincludes inner and outer cases joined by the airfoils. The leading andtrailing edges respectively extend in a radial direction from the innerand outer case a leading edge span and a trailing edge span. The airfoilextends in an axial direction an axial chord length between the leadingand trailing edges. The airfoils each have an aspect ratio of less than1.5, wherein the aspect ratio is an average of the sum of the leadingand trailing edge spans divided by the axial chord length.

In a further embodiment of any of the above, the low and high pressureturbines are configured to rotate in opposite directions.

In a further embodiment of any of the above, the first angle is greaterthan 20°.

In a further embodiment of any of the above, the array includes twentyor fewer airfoils.

In a further embodiment of any of the above, the gas turbine engineincludes a compressor section having a high pressure compressor and alow pressure compressor. A combustor is fluidly connected to thecompressor section, and a turbine section is fluidly connected to thecombustor. The turbine section includes a high pressure turbine thatprovides the first stage. A low pressure turbine provides the secondstage. A mid-turbine frame provides the frame positioned between thehigh pressure turbine and the low pressure turbine.

In another further embodiment of any of the foregoing gas turbine engineembodiments, a fan is fluidly connected to the compressor section.

In another further embodiment of any of the foregoing gas turbine engineembodiments, a geared architecture is interconnected between the fan andthe low pressure turbine.

In another further embodiment of any of the foregoing gas turbine engineembodiments, the gas turbine engine may be a high bypass geared aircraftengine having a bypass ratio of greater than about six (6).

In another further embodiment of any of the foregoing gas turbine engineembodiments, the gas turbine engine may include a low Fan Pressure Ratioof less than about 1.45.

In another further embodiment of any of the foregoing gas turbine engineembodiments, the low pressure turbine may have a pressure ratio that isgreater than about 5.

A gas turbine engine includes low and high pressure turbines that have arotational axis. A mid turbine frame is arranged axially between low andhigh pressure turbines. The mid turbine frame includes a circumferentialarray of airfoils. The mid turbine frame includes inner and outer casesjoined by the airfoils. Leading and trailing edges respectively extendin a radial direction from the inner and outer case a leading edge spanand a trailing edge span. The airfoil extends in an axial direction anaxial chord length between the leading and trailing edges. The airfoilseach have an aspect ratio range of greater than 1.0 to about 1.5,wherein the aspect ratio is an average of the sum of the leading andtrailing edge spans divided by the axial chord length.

In a further embodiment of any of the above, the airfoils each have acurvature provided equidistantly between pressure and suction sides andextend from the leading edge to the trailing edge at a midspan planealong the airfoil. A plane extends through the rotational axis andintersects the trailing edge and curvature. First and second lines arerespectively tangent to the curvature at the leading and trailing edges.A first angle is provided between the plane and the second line and asecond angle is provided between the second and first lines, wherein thesecond angle is greater than 10°.

In a further embodiment of any of the above, the midspan plane isoriented at a flow path angle relative to the rotational axis in therange of 20°-60°.

In a further embodiment of any of the above, the low and high pressureturbines are configured to rotate in opposite directions.

In a further embodiment of any of the above, the first angle is greaterthan 20°.

In a further embodiment of any of the above, the array includes twentyor fewer airfoils.

In a further embodiment of any of the above, the gas turbine engineincludes a compressor section fluidly having a high pressure compressorand a low pressure compressor. A combustor is fluidly connected to thecompressor section, and a turbine section is fluidly connected to thecombustor. The turbine section includes a high pressure turbine thatprovides the first stage. A low pressure turbine provides the secondstage. A mid-turbine frame provides the frame positioned between thehigh pressure turbine and the low pressure turbine.

In another further embodiment of any of the foregoing gas turbine engineembodiments, the gas turbine engine may be a high bypass geared aircraftengine having a bypass ratio of greater than about six (6).

In another further embodiment of any of the foregoing gas turbine engineembodiments, the gas turbine engine may include a low Fan Pressure Ratioof less than about 1.45.

In another further embodiment of any of the foregoing gas turbine engineembodiments, the low pressure turbine may have a pressure ratio that isgreater than about 5.

BRIEF DESCRIPTION OF THE DRAWINGS

The disclosure can be further understood by reference to the followingdetailed description when considered in connection with the accompanyingdrawings wherein:

FIG. 1 schematically illustrates an example gas turbine engine.

FIG. 2A is a front elevational view of an example mid turbine frameschematically depicting a bearing and oil and air sources.

FIG. 2B is a side perspective view of the mid turbine frame illustratedin FIG. 2A.

FIG. 3 is a cross-sectional view through a midspan plane of an airfoilshown in FIG. 4.

FIG. 4 is a schematic side view of an airfoil in the mid turbine frame.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. Alternative engines mightinclude an augmentor section (not shown) among other systems orfeatures. The fan section 22 drives air along a bypass flowpath whilethe compressor section 24 drives air along a core flowpath forcompression and communication into the combustor section 26 thenexpansion through the turbine section 28. Although depicted as aturbofan gas turbine engine in the disclosed non-limiting embodiment, itshould be understood that the concepts described herein are not limitedto use with turbofans as the teachings may be applied to other types ofturbine engines including three-spool architectures.

The engine 20 generally includes a low speed spool 30 and a high speedspool 32 mounted for rotation in opposite direction relative to oneanother about an engine central longitudinal axis A relative to anengine static structure 36 via several bearing systems 38. It should beunderstood that various bearing systems 38 at various locations mayalternatively or additionally be provided.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects a fan 42, a low pressure compressor 44 and a low pressureturbine 46. The inner shaft 40 is connected to the fan 42 directly orthrough a geared architecture 48 to drive the fan 42 at a lower speedthan the low speed spool 30. The high speed spool 32 includes an outershaft 50 that interconnects a high pressure compressor 52 and a highpressure turbine 54. A combustor 56 is arranged between the highpressure compressor 52 and the high pressure turbine 54. A mid-turbineframe 57 of the engine static structure 36 is arranged generally betweenthe high pressure turbine 54 and the low pressure turbine 46. Themid-turbine frame 57 supports one or more bearing systems 38 in theturbine section 28. The inner shaft 40 and the outer shaft 50 areconcentric and rotate via bearing systems 38 about the engine centrallongitudinal axis A, which is collinear with their longitudinal axes.

The core airflow is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over the high pressure turbine 54, themid-turbine frame 57, and low pressure turbine 46. The mid-turbine frame57 includes circumferential array of airfoils 59, which are arranged inthe core airflow path axially between the low and high pressure turbines46, 54. In one example, there are twenty or fewer airfoils arranged in asingle axial row circumferentially along the mid turbine frame flowpath. The turbines 46, 54 rotationally drive the respective low speedspool 30 and high speed spool 32 in response to the expansion.

The engine 20 in one example a high-bypass geared aircraft engine. In afurther example, the engine 20 bypass ratio is greater than about six(6), with an example embodiment being greater than ten (10), the gearedarchitecture 48 is an epicyclic gear train, such as a planetary gearsystem or other gear system, with a gear reduction ratio of greater thanabout 2.3 and the low pressure turbine 46 has a pressure ratio that isgreater than about 5. In one disclosed embodiment, the engine 20 bypassratio is greater than about ten (10:1), the fan diameter issignificantly larger than that of the low pressure compressor 44, andthe low pressure turbine 46 has a pressure ratio that is greater thanabout 5:1. Low pressure turbine 46 pressure ratio is pressure measuredprior to inlet of low pressure turbine 46 as related to the pressure atthe outlet of the low pressure turbine 46 prior to an exhaust nozzle.The geared architecture 48 may be an epicycle gear train, such as aplanetary gear system or other gear system, with a gear reduction ratioof greater than about 2.5:1. It should be understood, however, that theabove parameters are only exemplary of one embodiment of a gearedarchitecture engine and that the present invention is applicable toother gas turbine engines including direct drive turbofans.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft, withthe engine at its best fuel consumption—also known as “bucket cruiseThrust Specific Fuel Consumption (‘TSFC’)”—is the industry standardparameter of 1 bm of fuel being burned per hour divided by 1 bf ofthrust the engine produces at that minimum point. “Fan pressure ratio”is the pressure ratio across the fan blade alone, without a Fan ExitGuide Vane (“FEGV”) system. The fan pressure ratio as disclosed hereinaccording to one non-limiting embodiment is less than about 1.45. “Lowcorrected fan tip speed” is the actual fan tip speed in ft/sec dividedby an industry standard temperature correction of [(Tambient degR)/518.7)^0.5]. The “Low corrected fan tip speed,” as disclosed hereinaccording to one non-limiting embodiment, is less than about 1150ft/second.

Referring to FIGS. 2A and 2B, the mid turbine frame 57 includes innerand outer cases 58, 60 joined by the airfoils 59 to define a mid turbineframe flow path through which core airflow C passes. In one example, theairfoils 59 provide cavities 61 through which components and/or fluidsmay pass. For example, a structure support 87 may extend through thecavities 61 to support a bearing 86 arranged in a bearing compartment88. The bearing 86 is part of a bearing system 38, which may support thehigh speed spool 32. An oil source 90 may communicate oil to the bearingcompartment 88 through a cavity 88, and an air source 92 may supply airthrough a cavity 61 to buffer the bearing compartment 88.

Referring to FIGS. 3 and 4, with continuing reference to FIGS. 2A-2B,the airfoils 59 include circumferentially spaced pressure and suctionsides 66, 68 extending somewhat axially between leading and trailingedges 62, 64. Each airfoil 59 has a curvature 76 provided equidistantlybetween pressure and suction sides 66, 68. The airfoils 59 extend fromthe leading edge 62 to the trailing edge 64 along a midspan plane 75. Inone example, the midspan plane 75 is oriented at a flow path angle 84relative to the rotational axis A in the range of 20°-60°

The airfoils 59 have a camber that induces a turning airflow as the airpasses through the mid turbine frame 57 between the counter rotatingfirst and second stages, such as high and low pressure turbines 54, 46.It should be understood that the airfoils 59 may also be used betweenother sets of rotating stages. A plane P extends through the rotationalaxis A and intersects the trailing edge 64 and curvature 76. First andsecond lines 78, 80 are respectively tangent to the curvature 76 at theleading and trailing edges 62, 64. A first angle 81 is provided betweenthe plane P and the second line 80, and a second angle 82 is providedbetween the second and first lines 80, 78. In one example, the firstangle 81 is in a range of 0°-70°, and the second angle 82 is greaterthan 10°. It should be understood that the first angle may have othervalues outside the range and still fall within the scope of thisdisclosure.

Referring to FIG. 4, the leading and trailing edges 62, 64 respectivelyextend in a generally radial direction from the inner and outer cases58, 60 a leading edge span 72 and a trailing edge span 74. The airfoil59 extends in an axial direction an axial chord length 70 between theleading and trailing edges 62, 64. The airfoils 59 each have an aspectratio of less than 1.5, wherein the aspect ratio is an average of thesum of the leading and trailing edge spans 72, 74 divided by the axialchord length 70. In one example, the aspect ratio has a range of greaterthan 1.0 to about 1.5.

Although an example embodiment has been disclosed, a worker of ordinaryskill in this art would recognize that certain modifications would comewithin the scope of the claims. For that reason, the following claimsshould be studied to determine their true scope and content.

What is claimed is:
 1. A gas turbine engine comprising: a first stageand a second stage having a rotational axis; a frame arranged axiallybetween the first stage and the second stage, the frame including acircumferential array of airfoils, at least one of the airfoils having acurvature provided equidistantly between pressure and suction sides andextending from a leading edge to a trailing edge at a midspan planealong the airfoil, and an angle defined between first and second linesrespectively tangent to the intersection of the midspan plane and thecurvature at the airfoil leading and trailing edges, the angle beingequal to or greater than about 10°.
 2. The gas turbine engine accordingto claim 1, wherein the midspan plane is oriented at a flow path anglerelative to the rotational axis in a range of 20°-60°.
 3. The gasturbine engine according to claim 1, wherein the frame includes an innerand an outer case joined by the airfoils, the leading and trailing edgesrespectively extending in a generally radial direction from the innercase and the outer case a leading edge span and a trailing edge span,and the airfoil extends in an axial direction an axial chord lengthbetween the leading and trailing edges, the at least one of airfoilshaving an aspect ratio of less than 1.5, wherein the aspect ratio is anaverage of the sum of the leading and trailing edge spans divided by theaxial chord length.
 4. The gas turbine engine according to claim 1,wherein the first stage and the second stage are configured to rotate inopposite directions.
 5. The gas turbine engine according to claim 1,wherein a rotational axis plane extends through the rotational axis andintersects the trailing edge and the curvature, a first angle providedbetween the rotational axis plane and the second line, the angleproviding a second angle, and wherein the first angle is greater than20°.
 6. The gas turbine engine according to claim 1, wherein the arrayincludes twenty or fewer airfoils.
 7. The gas turbine engine accordingto claim 1, comprising: a compressor section comprising a high pressurecompressor and a low pressure compressor; a combustor fluidly connectedto the compressor section; a turbine section fluidly connected to thecombustor, the turbine section comprising: the first stage is a highpressure turbine; the second stage is a low pressure turbine; andwherein the frame is provided by a mid-turbine frame positioned betweenthe high pressure turbine and the low pressure turbine.
 8. The gasturbine engine according to claim 7, further comprising a fan fluidlyconnected to the compressor section.
 9. The gas turbine engine accordingto claim 8, comprising a geared architecture is interconnected betweenthe fan and the low pressure turbine.
 10. The gas turbine engineaccording to claim 8, wherein the gas turbine engine is a high bypassgeared aircraft engine having a bypass ratio of greater than about six(6).
 11. The gas turbine engine according to claim 8, wherein the gasturbine engine includes a Fan Pressure Ratio of less than about 1.45.12. The gas turbine engine according to claim 8, wherein the lowpressure turbine has a pressure ratio that is greater than about
 5. 13.A gas turbine engine comprising: a first stage and a second stage havinga rotational axis; a frame arranged axially between the first stage andthe second stage, the frame including a circumferential array ofairfoils, the frame includes inner and outer cases joined by theairfoils, leading and trailing edges respectively extending in agenerally radial direction from the inner and outer cases a leading edgespan and a trailing edge span, and the airfoil extends in an axialdirection an axial chord length between the leading and trailing edges,the airfoils each having an aspect ratio range having a lower limit ofgreater than 1.0 to an upper limit of about 1.5, wherein the aspectratio is an average of the sum of the leading and trailing edge spansdivided by the axial chord length.
 14. The gas turbine engine accordingto claim 13, the airfoils each having a curvature provided equidistantlybetween pressure and suction sides and extending from the leading edgeto the trailing edge at a midspan plane along the airfoil, a rotationalaxis plane extending through the rotational axis and intersecting thetrailing edge and the curvature, first and second lines respectivelytangent to the curvature at the leading and trailing edges, a firstangle provided between the rotational axis plane and the second line anda second angle provided between the second and first lines, wherein thesecond angle is greater than 10°.
 15. The gas turbine engine accordingto claim 14, wherein the midspan plane is oriented at a flow path anglerelative to the rotational axis in a range of 20°-60°.
 16. The gasturbine engine according to claim 13, wherein the first stage and thesecond stage are configured to rotate in opposite directions.
 17. Thegas turbine engine according to claim 14, wherein the first angle isgreater than 20°.
 18. The gas turbine engine according to claim 13,wherein the array includes twenty or fewer airfoils.
 19. The gas turbineengine according to claim 13, comprising: a compressor sectioncomprising a high pressure compressor and a low pressure compressor; acombustor fluidly connected to the compressor section; a turbine sectionfluidly connected to the combustor, the turbine section comprising: thefirst stage is a high pressure turbine; the second stage is a lowpressure turbine; and wherein the frame is provided by a mid-turbineframe positioned between the high pressure turbine and the low pressureturbine.
 20. The gas turbine engine according to claim 19, wherein thegas turbine engine includes a fan and is a high bypass geared aircraftengine having a bypass ratio of greater than about six (6).
 21. The gasturbine engine according to claim 20, wherein the gas turbine engineincludes a Fan Pressure Ratio of less than about 1.45.
 22. The gasturbine engine according to claim 19, wherein the low pressure turbinehas a pressure ratio that is greater than about
 5. 23. A mid turbineframe for a gas turbine engine, comprising: a frame of the mid turbineframe including a circumferential array of airfoils, at least one of theairfoils having a curvature provided equidistantly between pressure andsuction sides and extending from a leading edge to a trailing edge at amidspan plane along the airfoil, and an angle defined between first andsecond lines respectively tangent to the intersection of the midspanplane and the curvature at the airfoil leading and trailing edges, theangle being equal to or greater than about 10°.
 24. A turbine sectionfor a gas turbine engine, comprising: a first turbine and a secondturbine having a rotational axis; a mid turbine frame arranged axiallybetween the first turbine and the second turbine, a mid turbine frameincluding a circumferential array of airfoils, at least one of theairfoils having a curvature provided equidistantly between pressure andsuction sides and extending from a leading edge to a trailing edge at amidspan plane along the airfoil, and an angle defined between first andsecond lines respectively tangent to the intersection of the midspanplane and the curvature at the airfoil leading and trailing edges, theangle being equal to or greater than about 10°.
 25. The turbine sectionaccording to claim 24, wherein one of the first and second turbines is alow pressure turbine, the low pressure turbine has a pressure ratio thatis greater than about
 5. 26. The turbine section according to claim 24,wherein the array includes twenty or fewer airfoils.